The assignee of the present invention manufactures and deploys spacecraft for observation of the Earth and other celestial bodies, broadcast and communications purposes. To perform their mission, the payloads and solar panels of such spacecraft must be oriented and maintained in a particular orientation with respect to the Earth (or other celestial body) or with respect to the Earth and sun. For example, the general requirement for three-axis, body-stabilized spacecraft operating in geostationary orbit is to orient the spacecraft such that the payload is directed toward the Earth and the axis of rotation of the solar array is orthogonal to the Earth's equatorial plane. In such orientation, the solar arrays are enabled to rotate in such a manner to compensate for the spacecraft's motion with respect to the sun and thereby receive constant and continuous solar illumination to the photovoltaic cells on the array. Maintenance of this desired three-axis attitude is provided by way of attitude and rate sensors coupled to torque generators through an attitude determination and control subsystem, which includes an onboard computer referred to as the spacecraft control electronics. Attitude sensors may comprise celestial body observers, such as earth sensors, sun sensors and star trackers. Rate sensors may comprise such devices as digital integrating rate assemblies or gyros. Torque generators may comprise such devices as thrusters, magnetic torquers or momentum wheels.
The spacecraft design and operating methods must provide means to achieve the desired three-axis attitude upon the initiation of the mission and to reacquire this attitude following any planned or unplanned deviation from it. Because unplanned deviations from the desired three-axis attitude can result in service disruptions that are costly to the spacecraft operators and their customers, it is preferable that the spacecraft design and operating methods provide means to realign the spacecraft in the desired attitude in a prompt, rapid and reliable manner, with a minimal amount of operator intervention and at any time of the year and any time of day.
A number of known techniques for reacquiring Earth pointing attitude from a different known or unknown attitude have been described.
For example, U.S. Pat. No. 5,080,307 issued to Smay, et al., teaches a method of acquiring Earth-pointing attitude of a three-axis, body-stabilized spacecraft orbiting the Earth, including the steps of (1) aligning the roll axis to the spacecraft with the sun line; (2) orienting the spacecraft such that the angle formed between the yaw axis and the sun line is equal to the Earth-sun angle; (3) orienting the spacecraft such that the yaw axis is aligned with the center of the Earth; (4) rotating the spacecraft about its yaw axis until its pitch axis is oriented normal to the orbit plane. Disadvantages of this method include the necessity to perform four discrete maneuvers, and the fact that the method requires initially aligning the spacecraft roll axis with the sun line, both of which require expenditures of time and attitude control propellant which are minimized by the present invention.
U.S. Pat. No. 6,695,263, issued to Goodzeit, teaches a method for earth reacquisition by (1) determining three-axis inertial attitude by rotating the spacecraft slowly about its pitch axis while measuring star patterns; (2) adjusting the attitude to align the pitch axis with (parallel to) the Earth pole axis; (3) rotating the spacecraft about the pitch axis to establish communications with the ground. This method is disadvantageous compared to the present invention inasmuch as it requires over an hour to execute, requires expensive star trackers that are often not otherwise required for a spacecraft mission, and requires extensive intervention by spacecraft ground controllers.
U.S. Pat. No. 5,535,965 issued to Surauer, et al., teaches a sun and earth acquisition method avoiding dependence on gyros by use of wide field of view sun sensors and at least one earth sensor. Disadvantages of this method include a requirement for software to estimate spacecraft rotation rates from 2-axis position measurements resulting in very complicated algorithms. Additionally, certain spacecraft reorientations such as those needed during orbit transfer require additional earth sensors with a consequential added expense that is avoided by the teaching of the present invention. Finally, the method taught by Surauer, et al., results in acquiring the Earth in an earth sensor field of view but does not orient the spacecraft attitude angle about the Earth-to-spacecraft line. By contrast, the present invention uses a conventional hardware suite with gyros, to enable a reacquisition method that results in the spacecraft acquiring a desired 3-axis attitude.
U.S. Pat. No. 6,142,422 issued to Stoen, et al., teaches a method to orient the spacecraft along an inertial direction of interest by the steps of operating a sensor to provide an initial fix on the inertial direction of interest; repetitively determining a difference between a commanded quaternion and a quaternion estimated based on sensed angular rates; and selectively applying torques to the spacecraft so as to drive the difference towards zero such that a spacecraft vector is aligned with the inertia direction of interest, thereby orienting the spacecraft. Although the method disclosed by Stoen, et al., is useful in accomplishing the objectives of the present invention, Stoen does not teach a method of aligning a spacecraft in a three-axis attitude with respect to the Earth and the sun.
Still other methods of reacquiring Earth-pointing orientation are known that present similar difficulties with respect to time to reacquire, hardware cost and requirements for operator intervention.
Accordingly, it is an objective of the present invention to provide a method for aligning a spacecraft in a desired three-axis attitude with respect to both the sun and the Earth or other celestial body at any time of day in accordance with a method capable of autonomous or semi-autonomous execution.
It is a further objective of the present invention to provide a method for fast reacquisition of Earth-normal attitude using existing spacecraft resources of sensors, ephemeris and clock data, and onboard spacecraft computational capabilities.
Yet further objectives of the present invention are to minimize the burden on the spacecraft operator and eliminate dependency on telemetry and command coverage from the ground. The present invention does not require operator calculation of reorientation angles, ground generation of commands via manual commands or time-tagged commanding.